Combined cycle integrated combustor and nozzle system

ABSTRACT

An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of said DMRJ functioning as an aerodynamic choke.

CROSS REFERENCE TO RELATED APPLICATION(S)

N.A.

U.S. GOVERNMENT RIGHTS

N.A.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to a combined cycle engine system that combines aturbojet or other booster engine and a Dual Mode Ramjet (DMRJ) to permitefficient operation from takeoff to hypersonic speeds and moreparticularly from Mach 0 Sea Level to Mach 5+ at high altitude.

2. Description of the Related Art

A conventional DMRJ cannot produce thrust to accelerate itself tosupersonic speeds. The DMRJ must be boosted by some other propulsiveelement, such as a Turbojet Engine (TJ). When the booster engine is aturbojet (turbine) engine, the combined cycle engine is referred to as aTurbine Based Combined Cycle or TBCC engine. If the booster is a rocketengine, it is referred to as a Rocket Based Combined Cycle or RBCCengine. In prior art reusable hypersonic vehicle concepts, the turbineengines of a TBCC engine are expected to produce all of the thrust atsubsonic and low supersonic speeds. At some higher speed, the DMRJ isturned on to produce the required thrust and the TJ is turned off atnearly the same speed and taken out of the flow. The thrust of the TJengine as the sole propelling means during the bulk of the accelerationplaces a great demand on the TJ technologies. Prior art TBCC engineshave little or no DMRJ thrust contribution at speeds below Mach 3-4. Thehighest thrust requirement for these vehicles occurs during accelerationfrom subsonic to supersonic speeds. This so called “transonic” speed hasthe greatest drag to overcome. From the foregoing, it is seen that thegreater the thrust contribution from the DMRJ during acceleration theless demands are placed on the turbine engine.

Current turbine engine technology is suitable for speeds up to Mach 2.5.Above this speed, the air temperature becomes too high to permit highcompressor pressure ratio without exceeding the turbine entrancetemperature limits. This results in a reduction in engine airflow andthrust. A Lockheed SR-71 high-speed, high-altitude, reconnaissanceaircraft was able to fly at about Mach 3.25 by bypassing some of the airaround the final compressor stages in the Pratt & Whitney J-58 engine.This unloaded the compressor, reducing the combustion heat additionrequired upstream of the turbine. Such an engine cycle is referred to asa Turbo-Ramjet since most of the high Mach thrust is produced in anafterburner downstream of the turbine. This engine cycle is not apreferred cycle for a TBCC system which would need a third duct for DMRJoperation. Since the DMRJ duct must operate at speeds beyond what the TJcan stand, the DMRJ air flow must bypass the TJ completely. Oneimportant issue with the prior art TBCC engines is the requirement forthe Turbine engine to operate to Mach 4 or higher. This places a largetechnical hurtle to develop a pure Turbojet that can operate with highthrust at Mach 4 or higher. The thrust a TJ can produce as a function ofMach number is dependent on the technologies applied. For Mach 4operation at high thrust, advanced high strength high temperaturematerials are needed that are not currently available.

When not operating, the DMRJ flowpath increases vehicle drag if airflows through the duct or around it. At speeds below typically Mach 5,the TBCC nozzle is over-expanded (too large) which reduces the netthrust. Increasing the size of the turbine to produce sufficient thrustto overcome the vehicle and the non-operating DMRJ engine drag hassevere mission consequences due to greater vehicle empty weight andreduced available fuel volume. United States patent applicationpublication number US 2006/0107648 A1 by Bulman et al. discloses a TBCChaving an integrated inlet that manages the flow of air to both the TJand DMRJ elements. The US 2006/0107648 A1 patent application isincorporated by reference in its entirety herein.

It is known that the thrust in a DMRJ at low speeds is limited due tolow ram pressure and premature thermal choking of the combustor. Weaddress each of the limiting factors on low Mach TBCC thrust:

Subsonic to Low Supersonic Thrust—As a ram compression cycle, the DMRJhas little thrust potential at low speeds. For typical TBCC poweredhypersonic vehicles the drag at transonic speeds (Mach 0.8-1.3) isusually more than the turbine engine can produce. Additional thrust isneeded but just installing a larger turbojet engine is unattractive in aweight and volume sensitive hypersonic vehicle.

Low Supersonic to Mach 4 Thrust—Prior art DMRJs have a Scramjetdiverging combustor and an isolator to allow operation with a thermalthroat. These engines are usually not capable of producing useful thrustmuch below about Mach 4. A first factor is a low inlet/isolator pressurerise capability at low speeds. A second factor is that at low speeds andtypical combustor area ratios, the pressure rise for a given fuelequivalence ratio increases at low supersonic speeds. FIG. 1 shows thetemperature rise to thermally choke a DMRJ as a function of speed andcombustor Area Ratio (AR). A typical prior art DMRJ has a small AR, onthe order of 2 (Reference line 10), suitable for the higher speeds. Theengine thrust is directly related to the temperature rise. If too muchheat is added with a low AR combustor, the combustor pressure willexceed the inlet capability and the inlet will unstart. This miss-matchbetween available pressure and combustor back pressure preventspractical thrust from a prior art DMRJ. Below Mach 4, a high area ratiocombustor improves this situation (Reference line 12), but isinefficient for high speed operation since it increases the wetted areaexposed to the high speed, high enthalpy, flow—increasing engine weightand heat load while reducing high speed thrust.

TBCC Combined Thrust—Typical prior art hypersonic cruise vehicles haveas large a nozzle exit area as practical since at cruise speed theexhaust is under-expanded and thrust and Isp increase with largernozzles. At low speeds, the exhaust is then over-expanded and thrust andIsp are lower than they would be with a smaller nozzle. One solution tothis problem, creating a larger volume of gas to help fill andpressurize the nozzle, is disclosed in U.S. Pat. No. 6,568,171 toBulman. U.S. Pat. No. 6,568,171 is incorporated by reference in itsentirety herein.

BRIEF SUMMARY OF THE INVENTION

Accordingly, an object of the invention is to provide an engine thatoperates and produces the entire required vehicle thrust below Mach 4. Afeature of this object is that a Hypersonic combined cycle vehiclebecome more near term, thereby saving further vehicle and enginedevelopment costs. One such engine is a turbine based combined cycle(TBCC) engine having both a turbine and a dual mode ramjet (DMRJ). Theturbine and the DMRJ are integrated to provide effective thrust fromMach 0 to in excess of Mach 4.

As the TBCC accelerates the vehicle from Mach 0 to in excess of Mach 4,from Mach 0 to about Mach 2 incoming air delivered to the DMRJ isaccelerated by primary ejector thrusters that may receive oxidizer fromeither on-board oxidizer tanks or from turbine compressor discharge air.As the TBCC further accelerates the vehicle from about Mach 0 to inexcess of Mach 4 exhaust from the turbine and exhaust from the DMRJ arecombined in a common nozzle disposed downstream of a combustor portionof said DMRJ functioning as an aerodynamic choke.

The details of one or more embodiments of the invention are set forth inthe accompanying drawings and the description below. Other features,objects and advantages of the invention will be apparent from thedescription and drawings, and from the claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates the allowed combustor temperature rise as a functionof both flight Mach number and combustor area ratio.

FIG. 2 illustrates a TBCC engine configured for acceleration fromstationary to low supersonic speed in accordance with the invention.

FIG. 3 illustrates an improvement in thrust and Isp as a function of ERJaugmentation.

FIG. 4 illustrates a TBCC engine configured for acceleration from lowsupersonic to about Mach 4.

FIG. 5 illustrates how an underexpanded plume overshoots its ideal flowarea.

Like reference numbers and designations in the various drawingsindicated like elements.

DETAILED DESCRIPTION

This invention improves the performance of a combined cycle engineduring the critical acceleration from takeoff to ramjet takeover andincludes two elements that significantly increase the combined thrust ofa booster and DMRJ components when operating in parallel (Mach 0-3+). Atspeeds below where a ramjet is normally capable of operating, turbinecompressor bleed air or onboard oxidizers are used to drive smallprimary ejector thrusters to induce airflow into a dual mode ramjetflowpath and produce thrust as a well known Ejector Ramjet (ERJ). Inaddition, I integrate the DMRJ and booster flows in a synergistic way toincrease the overall engine thrust. In this integration, both enginesexhausts are merged into a common nozzle. Underexpanded turbine exhaustis used to create an aerodynamic choke for the DMRJ. This techniquesolves a critical problem with a low speed DMRJ by providing a largercombustor area and increases the thrust without unstarting the inlet.This allows higher combustion temperature in the DMRJ flow whichincreases the volume of gases in the nozzle increasing the pressure andthrust while reducing overexpansion losses. In this disclosure, Idiscuss the combustor and nozzle integration. The high Mach Turbojetengine thrust and technology requirements can be reduced by using a DMRJengine integration that provides high thrust at low speeds to contributeto the total vehicle thrust.

This invention is an improved combined cycle propulsion system that morecompletely and efficiently combines the operation of a booster and DualMode Ramjet engines. Parallel operation of the DMRJ with the boosterengine is enhanced by employing Ejector pumping at low speeds and higherDMRJ combustor area ratio at mid speeds via an aerodynamic choking withthe booster exhaust. The combined nozzle flows are more efficient thanseparate nozzles.

During transition from subsonic to low supersonic thrust, to contributethrust when it is needed most, we convert the DMRJ into an EjectorRamjet (ERJ). With reference to FIG. 2, the ejector primary thrusters 34are provided with oxidizer from one of two sources. On board oxidizertanks 31 are one source. This combined cycle configuration is called aTurbine/Rocket Based Combined Cycle (T/RBCC) engine. A second source isto take a small amount of the compressor discharge air (CompressorBleed) (˜10%) from a turbojet 30. The oxidizer is fed by supply lines 30a and 31 a. The ejector primaries are located in the trailing edge ofstruts in the DMRJ. FIG. 2 illustrates both concepts.

A TBCC engine 20 has a forward facing air inlet 22 with an internalstreamline 24 that divides the incoming air into a first airflow portion26 and a remainder airflow portion 28. The first airflow portion, whichat low speeds, nominally comprises 80% of the volume of air, feeds aturbine engine 30. The remainder airflow portion 28 enters the DMRJcombustor 32 which at low speeds (below about Mach 2) is used as anEjector/Mixer. The induced secondary air is accelerated by a pluralityof primary ejectors 34. The primary ejectors work by the well knownprinciples of ejector pumping through viscous coupling between theprimary and secondary flows.

Primary ejectors 34 are disposed on the trailing edges of struts 36.Fuel injectors (not shown) are positioned to inject fuel into the DMJRin the vicinity of an ejector mixing region 38. Typically, these fuelinjectors are also located on the trailing edges of struts 36 or alongwalls 40 of the ejector mixing region 38. The fuel/air mixture isignited by a suitable pilot (not shown) and combusted in ejectorcombustor 42.

Removing the bleed air from the turbine engine 30 causes a loss inturbine engine thrust but the ejector pumping process induces aboutthree times that much additional airflow through the DMRJ flowpath atlow speeds. The subsequent combustion of this additional air results ina 10-20% increase in net engine thrust. The Isp drops a little due tothe low pressure in the DMRJ but the net effective Isp (Ieff) isincreased. Ieff=(F-D)/(fuel flowrate). FIG. 3 shows how we expect thethrust and Isp to vary with the ERJ Augmentation.

During acceleration from low supersonic (about Mach 2) to about Mach 4,in order to increase the DMRJ thrust given its allotted airflow, we needto burn more fuel without causing the inlet to unstart. This requires alarger combustor area to avoid premature thermal choking, as previouslydiscussed. As shown in FIG. 4, in our integrated engine concept, weemploy a nozzle design that uses the turbine exhaust plume to create anaerodynamic blockage or secondary throat downstream of a low AR DMRJcombustor 32 exit within the DMRJ nozzle. This aerodynamic choke locatedwell aft in the common nozzle 42 creates additional combustor flow area(and distance) allowing more fuel to be burned without an inlet unstart.Since more fuel can be burned than without this technique, significantlyhigher low speed DMRJ thrust results without adverse consequences athigher speeds.

FIG. 4 illustrates the TBCC engine 20 configured for continuedacceleration from low supersonic (Mach ˜2) to about Mach 4. Downstreamof DMRJ combustor 32 is a common nozzle 42. Exhaust 46 from turbineengine 30 is directed alongside an exterior wall of the common nozzle42. A nozzle flap 48 opens during this phase of the mission enablingexhaust 46 to flow into the common nozzle 42 forming an aerodynamicchoke 50.

Since the turbine exhaust 46 is injected into the common nozzle 42, theflow area remaining for the DMRJ flow is reduced. By selecting thelocation of the turbine engine nozzle and its degree of expansion, wecan create an aerodynamic throat 50 at a larger area than would bepossible without this exhaust interaction. A larger throat area permitsgreater temperature increase at low speeds. Contributing to this effectis the use of an underexpanded turbine nozzle exit pressure. When theturbine nozzle flow is underexpanded, its exit pressure is greater thanthe prevailing pressure in the DMRJ nozzle 42. As the turbine flow exitsits nozzle, it will expand further than even its ideal equilibrium flowarea. This process is similar to the Fabri blockage seen in an EjectorRamjet Engine. FIG. 5 shows how an underexpanded plume overshoots itsideal flow area. This plume overshoot allows the turbine engine nozzleto be located further back in the DMRJ nozzle further increasing theeffective DMRJ combustor area ratio and thrust. This technique is animprovement over using a large complex and heavy variable geometrynozzle to achieve the same effect.

In FIG. 5:

A_(e)=Turbine Nozzle Exit Area.

A_(P)=Actual Turbine Exhaust Plume Area.

A_(pi)=Ideal Turbine Exhaust Plume Area (P=P_(s)).

K=Plume Overshoot Factor.

P_(C)=Turbine Exhaust Total Pressure.

P_(e)=Turbine Nozzle Exit Pressure.

P_(S)=Effective plume pressure.

γ_(r)=Turbine Exhaust Specific Heat Ratio.

When at supersonic speeds (less than Mach 5), the parallel operation ofthe DMRJ combustor with large heat addition generates a larger volume ofgas that helps fill and pressurize the common nozzle 42. In a typicalMach 2.5 case, the nozzle exit pressure increases from about ⅓ ofambient to about 75% ambient pressure by the addition of the DMRJ flow.As disclosed in U.S. Pat. No. 6,568,171, the thrust of both streams isincreased. The combined thrust is up to about 50% higher than theturbine alone if it had to fill the nozzle on its own.

One or more embodiments of the present invention have been described.Nevertheless, it will be understood that various modifications may bemade without departing from the spirit and scope of the invention. Forexample, a rocket or other low speed accelerator can be used in place ofthe turbine engine without deviating from the principles of theinvention. Accordingly, other embodiments are within the scope of thefollowing claims.

1. A combined cycle engine, comprising: a booster; and a dual moderamjet (DMRJ), wherein said booster and said DMRJ are integrated toprovide effective thrust from Mach 0 to in excess of Mach
 4. 2. Theengine of claim 1 wherein said booster and said DMRJ have a common airinlet.
 3. The engine of claim 2 wherein said common air inlet includes astreamline effective to divert a portion of incoming air to said DMRJand a remaining portion of said incoming air to said turbine.
 4. Theengine of claim 3 wherein said DMRJ includes a plurality of strutshaving primary ejector thrusters affixed to a trailing edge thereof. 5.The engine of claim 4 further including on-board oxidizer tanks coupledto said primary ejector thrusters.
 6. The engine of claim 4 wherein anexhaust portion of said booster is effectively coupled to said DMRJ toprovide turbine compressor discharge air to said primary ejectorthrusters.
 7. The engine of claim 4 wherein exhaust from said boosterand exhaust from said DMRJ are selectively combined in a common nozzlelocated downstream of a DMRJ combustor.
 8. The engine of claim 7 whereinsaid common nozzle has a larger cross-sectional area than said DMRJcombustor.
 9. The engine of claim 8 wherein a nozzle flap is effectiveto control combination of said booster exhaust and said DMRJ exhaust.10. The engine of claim 1 wherein said DMRJ includes a plurality ofstruts having primary ejector thrusters affixed to a trailing edgethereof.
 11. The engine of claim 10 further including on-board oxidizertanks coupled to said primary ejector thrusters.
 12. The engine of claim10 wherein an exhaust portion of said booster is effectively coupled tosaid DMRJ to provide turbine compressor discharge air to said primaryejector thrusters.
 13. The engine of claim 10 wherein exhaust from saidbooster and exhaust from said DMRJ are selectively combined in a commonnozzle located downstream of a DMRJ combustor.
 14. The engine of claim13 wherein said common nozzle has a larger cross-sectional area thansaid DMRJ combustor.
 15. The engine of claim 14 wherein a nozzle flap iseffective to control combination of said booster exhaust and said DMRJexhaust.
 16. The engine of claim 1 wherein exhaust from said booster andexhaust from said DMRJ are selectively combined in a common nozzlelocated downstream of a DMRJ combustor.
 17. The engine of claim 16wherein said common nozzle has a larger cross-sectional area than saidDMRJ combustor.
 18. The engine of claim 17 wherein a nozzle flap iseffective to control combination of said turbine exhaust and said DMRJexhaust.
 19. A method for accelerating a vehicle from Mach 0 to inexcess of Mach 4 comprising the steps of: a). providing said vehiclewith a combined cycle engine that includes a booster having integratedair flow with a dual mode ramjet (DMRJ) said turbine and said DMRJhaving a common air inlet wherein said common air inlet includes astreamline effective to divert a first portion of incoming air to saidDMRJ and a remaining portion of said incoming air to said booster, saidDMRJ having a plurality of struts having primary ejector thrustersaffixed to a trailing edge thereof; b). from Mach 0 to about Mach 2,accelerating said first portion of incoming air in said primary ejectorthrusters; and c). from about Mach 0 to in excess of Mach 4, combiningexhaust from said booster and exhaust from said DMRJ in a common nozzledisposed downstream of a combustor portion of said DMRJ.
 20. The methodof claim 19 wherein, in step (b), oxidizer to power said primary ejectorthrusters is delivered from either on-board oxidizer tanks or fromcompressor discharge air of a turbine engine.
 21. The method of claim 20wherein about 10%, by volume, of said compressor discharge air isdelivered to said primary ejector thrusters.
 22. The method of claim 21including providing said common nozzle with both a cross-sectional arealarger than said combustor portion of said DMRJ and with an aerodynamicchoke.
 23. The method of claim 22 wherein said aerodynamic choke isactuated by combining exhaust from said turbine with exhaust from saidcombustor portion of said DMRJ and said choke is disabled by separatingexhaust from said turbine from exhaust from said combustor portion ofsaid DMRJ.
 24. The method of claim 19 including providing said commonnozzle with both a cross-sectional area larger than said combustorportion of said DMRJ and with an aerodynamic choke.
 25. The method ofclaim 24 wherein said aerodynamic choke is actuated by combining exhaustfrom said booster with exhaust from said combustor portion of said DMRJand said choke is disabled by decreasing the volume of exhaust from saidbooster combined with exhaust from said combustor portion of said DMRJ.26. The method of claim 25 wherein a nozzle flap is opened to actuatesaid aerodynamic choke and said nozzle flap is closed to disable saidaerodynamic choke.